Gas turbine engine composite vane assembly and method for making same

ABSTRACT

A gas turbine engine composite vane assembly and method for making same are disclosed. The method includes providing at least two gas turbine engine airfoil composite preform components. The airfoil composite preform components are interlocked with a first locking component so that mating faces of the airfoil composite preform components face each other. A filler material is inserted between the mating surfaces of the airfoil composite preform components.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to and the benefit of U.S. ProvisionalPatent Application No. 61/774,987, filed 8 Mar. 2013, the disclosure ofwhich is now expressly incorporated herein by reference.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

Embodiments of the present disclosure were made with Unites Statesgovernment support under Contract No. FA8650-07-C-2803. The governmentmay have certain rights.

TECHNICAL FIELD

The present application relates to gas turbine engine ceramic matrixcomposite vane assemblies and methods for forming same, and moreparticularly to multiple-component gas turbine engine CMC assemblies andmethods for forming same.

BACKGROUND

Gas turbine engine ceramic matrix composite vane assemblies remain anarea of interest. Some existing systems have various shortcomings,drawbacks, and disadvantages relative to certain applications.Accordingly, there remains a need for further contributions in this areaof technology.

SUMMARY

One embodiment of the present disclosure is a unique method for forminga gas turbine engine ceramic matrix composite vane assembly in which,among other things, a locking component and/or mat filler material maybe provided in a joint portion of the assembly. Other embodimentsinclude unique methods, systems, devices, and apparatus for forming aCMC assembly. Further embodiments, forms, objects, aspects, benefits,features, and advantages of the present application shall becomeapparent from the description and figures provided herewith.

BRIEF DESCRIPTION OF THE FIGURES

Features of the application will be better understood from the followingdetailed description when considered in reference to the accompanyingdrawings, in which:

FIG. 1 is an exploded perspective view of a ceramic matrix composite(CMC) assembly according to an embodiment.

FIG. 2 is an end elevational view of the CMC assembly of FIG. 1 taken atelevation 2-2 of FIG. 1.

FIG. 3 is a side elevational view of the CMC assembly of FIG. 1 taken atelevation 3-3 of FIG. 1.

FIG. 4 is a flowchart depicting steps according to an embodiment of aprocess for forming a CMC assembly.

FIG. 5 is an exploded perspective view of a ceramic matrix composite(CMC) assembly according to another embodiment.

FIG. 6 is a perspective view of the CMC assembly of FIG. 4.

FIG. 7 is a flowchart depicting steps according to an embodiment of aprocess for forming a CMC assembly.

DETAILED DESCRIPTION OF REPRESENTATIVE EMBODIMENTS

While the present invention can take many different forms, for thepurpose of promoting an understanding of the principles of theinvention, reference will now be made to the embodiments illustrated inthe drawings and specific language will be used to describe the same. Itwill nevertheless be understood that no limitation of the scope of theinvention is thereby intended. Any alterations and further modificationsof the described embodiments, and any further applications of theprinciples of the invention as described herein, are contemplated aswould normally occur to one skilled in the art to which the inventionrelates.

FIG. 1 shows an exploded perspective view of a ceramic matrix composite(CMC) assembly 10 according to an embodiment. In the FIG. 1 embodiment,the CMC assembly 10 comprises components suitable for use in a gasturbine engine, although the CMC assembly 10 is not limited as such andother embodiments are contemplated herein. For example, the CMC assembly10 can comprise vanes and endwalls, cases with shields, or static flowcomponents. In one form, the CMC assembly 10 can comprise ceramic matrixcomponents for use in the hot section of a nuclear reactor.

Referring to FIG. 1, the CMC assembly 10 includes an upstanding airfoil12, an endwall 14, locking components 18, and a mat filler material 20.As will be described in greater detail below, the locking components 18can lock together the airfoil 12 and endwall 14, and the mat fillermaterial 20 can serve as a bond initiator and/or joint filler toaccommodate for example misalignment and/or tolerance errors at themating faces of the airfoil 12 and the endwall 14 or at the mating facesof the locking component 18 and the airfoil 12 and/or the endwall 14.

The airfoil 12 and endwall 14 can be preform partially-rigidized orun-rigidized components. The components can be fabricated of woven ornon-woven fiber. The fibers can be arranged and fixed by any suitabletechnique for example as by lay-up of fabrics, filament winding,braiding, knotting, or any combination of these. Further, the componentscan be partially or fully densified, or partially or fully infiltratedso as to fill in one or more gaps between fibers of matrix material. Thecomponents can also be of near-net shape and/or machined and/or undergofurther treatments such as coating or impregnation of the matrixmaterial, in order to, for example, provide features that constrain theCMC assembly 10 when used in conjunction with one or more lockingcomponents 18. It will be appreciated that the CMC assembly 10 can beconstructed of preform components having different configurations asnecessary or desired for a particular application. In the FIG. 1embodiment, for example, the airfoil 12 and endwall 14 each comprise apreform, and are trimmed and machined.

The airfoil 12 has a pressure side 22 and a suction side 24 that definea hollow 28 therebetween. A pair of airfoil connecting tabs 32, 34 canbe formed and/or machined at one end 30 of the airfoil 12 to extend inthe spanwise direction of the airfoil 12. The airfoil connecting tabs32, 34 each have through holes 36 that are sized to receive the lockingcomponents 18, as will be described in greater detail below. One airfoilconnecting tab 32 projects outwardly from and has substantially the samecontour as the pressure side 22 of the airfoil 12. The other airfoilconnecting tab 34 projects outwardly from and has substantially the samecontour as the suction side 24 of the airfoil 12. At opposite ends ofthe connecting tabs 32, 34 in the chordwise direction of the airfoil 12,the end 30 of the airfoil 12 can be configured to form a pair of seatportions 46, 48 that interface with the endwall 14 and/or the mat fillermaterial 20.

The endwall 14 includes a platform portion 52 having an opening 56 thatis sized to receive therethrough the airfoil connecting tabs 32, 34. Onopposite sides of the opening 56, a pair of endwall connecting tabs 62,64 can be formed and/or machined in the endwall 14 to correspond to thesuction side 24 and pressure side 22 of the airfoil 12 in a state wherethe endwall 14 is assembled to the airfoil 12 in the FIG. 1 embodiment.As illustrated in FIG. 1, the endwall connecting tabs 62, 64 cancorrespond substantially in size and shape to the airfoil connectingtabs 32, 34 of the airfoil 12. The endwall connecting tabs 62, 64 canproject in a manner that is upstanding relative to the platform portion52 of the endwall 14 and substantially parallel relative to the airfoilconnecting tabs 32, 34 of the airfoil 12. Referring to FIGS. 1 and 3, atopposite ends of the opening 56 corresponding to opposite ends of theairfoil connecting tabs 32, 34 in the spanwise direction of the airfoil12, the platform portion 52 of the endwall 14 has cooperating surfaces66, 68 (underside of the endwall 14 as shown in FIGS. 1 and 3) that restupon or over the seat portions 46, 48, respectively, at the end 30 ofthe airfoil 12. The cooperative relationship between the cooperatingsurfaces 66, 68 and the seat portions 46, 48 resists or prevents theendwall 14 from translating down the airfoil 12, as would be theinclination in the instance where for example greater pressure isapplied to the top side (as shown in FIG. 1) of the endwall 14 than theunderside thereof. Like the airfoil connecting tabs 32, 34, the endwallconnecting tabs 62, 64 each have through holes 76 that are sized toreceive the locking components 18.

As illustrated in FIG. 2, in an assembled state the locking components18 pass through the through holes 36, 76 in the airfoil 12 and endwall14. When fitted in the through holes 36, 76, the locking components 18lock the airfoil 12 into place relative to the endwall 14, preventingwithdrawal of the airfoil 12 from the endwall 14 in the direction fromwhich the airfoil 12 was inserted into the endwall 14, i.e. in thespanwise direction of the airfoil 12. The locking components 18 can takethe form of locking pins, although the locking components 18 are notlimited as such, and other embodiments are contemplated. For example,the locking components 18 can be sized to fit into a feature such as athrough-hole or cavity in and/or between one or more of the preformairfoil and endwall components. With respect to the FIG. 1 embodiment,as shown in FIG. 2, the locking components 18 can be substantially thesame length as the combined thickness of the endwall connecting tab 62,the airfoil connecting tab 32, and the mat filler material 20therebetween. The locking components 18 can be fabricated of braidedwoven or non-woven fiber. Further, the locking components 18 can bepartially or fully densified. The locking components 18 can also be ofnear-net shape and/or machined in order to, for example, providefeatures that facilitate constraining the preform components of the CMCassembly 10. In the FIG. 1 embodiment, for example, the lockingcomponents 18 each comprise a preform, and are trimmed and machined.

Referring now to FIGS. 2 and 3, a mat filler material 20 can be providedbetween the mating faces of the airfoil 12 and the endwall 14. Forexample, as shown in FIG. 2, the mat filler material 20 can be providedbetween the endwall connecting tab 62 of the endwall 14 and the airfoilconnecting tab 32 of the airfoil 12. As shown in FIG. 3, the mat fillermaterial 20 can be provided between the cooperating surface 66 of theendwall 14 and the seat portion 46 of the airfoil 12. The mat fillermaterial 20 can serve for example as a joint filler material and/or bondinitiator during for example chemical vapor infiltration (CVI)processing. The mat filler material 20 can comprise for example pre-cutpieces that substantially match for example the dimensions of the matingfaces of the airfoil 12 and the endwall 14. The mat filler material 20can be utilized for example to close unsuitable gaps between machined ornear-net partially-rigidized components. The mat filler material 20 canbe of non-woven material. In one example, the mat filler material 20 canbe in its raw, unprocessed state. As will be appreciated, the mat fillermaterial 20 can be of any thickness, or a varying thickness, that isnecessary or desired for a particular application.

FIG. 4 is a flowchart depicting steps of a process for forming a CMCassembly according to an embodiment. The airfoil 12 and endwall 14preform components can be provided in their un-rigidized orpartially-rigidized states (Step 80). The formed or machined airfoilconnecting tabs 32, 34 can be pushed through the opening 56 in theendwall 14 so that the airfoil connecting tabs 32, 34 are alongside theendwall connecting tabs 62, 64, as shown for example in FIG. 2, and thecooperating surfaces 66, 68 of the endwall 14 rest over or upon the seatportions 46, 48 of the airfoil 12, as shown for example in FIG. 3. Themat filler material 20 can be placed between the mating faces of theairfoil 12 and endwall 14, for example between the cooperating surface66 and the seat portion 46, either before or after the airfoilconnecting tabs 32, 34 are pushed through the opening 56 (Step 82). Thelocking components 18 can be inserted into the through holes 36 and 76in the respective airfoil connecting tab 32 and endwall connecting tab62, as shown for example in FIG. 2, and into the through holes 36 and 76in the respective airfoil connecting tab 34 and endwall connecting tab64, to lock the airfoil 12 to the endwall 14 (Step 84). The mat fillermaterial 20 can be placed between the mating faces of the airfoil 12 andthe endwall 14, for example between the airfoil connecting tab 32 andthe endwall connecting tab 62 as shown in FIG. 2, either before or afterthe airfoil connecting tabs 32, 34 are pushed through the opening 56, orbefore or after the locking components 18 are inserted into the throughholes 36 and 76. The assembled airfoil 12, endwall 14, lockingcomponents 18, and mat filler material 20 can be rigidized using a vaporinfiltration process (Step 86). As will be appreciated, any suitableprocess can be used for rigidizing the components, including for examplechemical vapor infiltration, slurry/melt infiltration, polymerinfiltration process, combined infiltration processes, to name just afew.

FIGS. 5 and 6 show a CMC assembly 110 according to an anotherembodiment. Like the FIG. 1 CMC assembly 10, the CMC assembly 110includes an airfoil 112 having airfoil connecting tabs 132, 134 that fitthrough an opening 156 of the endwall 114, and the endwall 114 restsover or upon seat portions 146, 148 of the airfoil 112. Further, likethe FIG. 1 embodiment, the FIG. 5 embodiment can have a mat fillermaterial 120 provided between mating faces of the airfoil 112 andendwall 114, for example, between the cooperating surfaces 166 and 168of the endwall 114 (underside of the endwall 114 as shown in FIGS. 4 and5) and the seat portions 146 and 148 of the airfoil 112.

The CMC assembly 110 has locking components 180, 182 that lock theairfoil 112 and endwall 114 together in a manner different from that ofthe locking components 18 of the FIG. 1 embodiment. The lockingcomponents 180 can take the form of substantially rectangular shapelocking members. As shown in FIG. 5, the rectangular shape lockingmembers 180 can be sized to fit into corresponding pre-machined orpre-formed substantially rectangular shape slots 186, 188 in therespective airfoil connecting tabs 132, 134. The slots 186, 188 can bepositioned immediately above the endwall 114 in a state where theendwall 114 is assembled to the airfoil 112 in the FIG. 5 embodiment.When the rectangular shape locking members 180 are fitted in the slots186, 188, the rear portions of the rectangular shape locking members 180rest over or upon the endwall 114 to lock the airfoil 112 into placerelative to the endwall 114. As shown in FIG. 6, the endwall 114 can beinterlocked between the rectangular shaped locking members 180 and theseat portions 146 and 148 of the airfoil 112, which prevents withdrawalof the airfoil 112 from the endwall 114 in the direction from whichairfoil 112 was inserted into the endwall 114, i.e. in the spanwisedirection of the airfoil 112.

The locking components 182 can take the form of locking pin members.Referring again to FIG. 5, the locking pin members 182 can be sized tofit into corresponding through holes 192 in the rectangular shapelocking members 180 and through holes 196 in the endwall 114. Thelocking pin members 182 can be substantially the same length as thecombined thickness of the endwall 114 and the rectangular shape lockingmember 180. The locking pin members 182 can serve to maintain structuralintegrity of the CMC assembly 110 by for example preventing withdrawalof the rectangular shape locking members 180 from the slots 186, 188during subsequent processing or at any time during the life of thecomponents should for example the bond between an airfoil connecting tab132, 134 and its corresponding rectangular shape locking member 180fail.

FIG. 7 is a flowchart depicting steps of a process for making a CMCassembly according to an embodiment. The airfoil 112 and endwall 114preform components can be provided in their un-rigidized orpartially-rigidized states (Step 200). Next, the formed or machinedairfoil connecting tabs 132, 134 can be pushed through the opening 156in the endwall 114 so that the airfoil connecting tabs 132, 134 projectthrough the opening 156 and the cooperating surfaces 166, 168 of theendwall 114 rest over or upon the seat portions 146, 148 of the airfoil112, as shown for example in FIG. 6. The mat filler material 120 can beplaced between the mating faces of the airfoil 112 and endwall 114, forexample between the cooperating surface 166 and the seat portion 146,either before or after the airfoil connecting tabs 132, 134 are pushedthrough the opening 156 (Step 202). The first locking components, thatis the rectangular shape locking members 180, can be inserted into theslots 186, 188 in the respective airfoil connecting tabs 132, 134, asshown for example in FIG. 6, to lock the airfoil 112 to the endwall 114(Step 204). In the illustrated embodiment, the inward facing surfaces ofthe rectangular shape locking members 180 are substantially alignedwith, that is flush with, the corresponding inward facing surfaces ofthe airfoil connecting tabs 132, 134. The second locking components,that is, the locking pin members 182, can be inserted into thecorresponding through holes 192 in the rectangular shape locking members180 and the through holes 196 in the endwall 114, as shown in FIG. 6(Step 206). The assembled airfoil 112, endwall 114, locking components180, 182, and mat filler material 120 can be rigidized using a vaporinfiltration process (Step 208). As will be appreciated, any suitableprocess can be used for rigidizing the components, including for examplechemical vapor infiltration, slurry/melt infiltration, polymerinfiltration process, combined infiltration processes, to name just afew.

Any theory, mechanism of operation, proof, or finding stated herein ismeant to further enhance understanding of embodiment of the presentinvention and is not intended to make the present invention in any waydependent upon such theory, mechanism of operation, proof, or finding.In reading the claims, it is intended that when words such as “a,” “an,”“at least one,” or “at least one portion” are used there is no intentionto limit the claim to only one item unless specifically stated to thecontrary in the claim. Further, when the language “at least a portion”and/or “a portion” is used the item can include a portion and/or theentire item unless specifically stated to the contrary.

While embodiments of the invention have been illustrated and describedin detail in the drawings and foregoing description, the same is to beconsidered as illustrative and not restrictive in character, it beingunderstood that only the selected embodiments have been shown anddescribed and that all changes, modifications and equivalents that comewithin the spirit of the invention as defined herein of by any of thefollowing claims are desired to be protected. It should also beunderstood that while the use of words such as preferable, preferably,preferred or more preferred utilized in the description above indicatethat the feature so described may be more desirable, it nonetheless maynot be necessary and embodiments lacking the same may be contemplated aswithin the scope of the invention, the scope being defined by the claimsthat follow.

What is claimed is:
 1. A method for forming a gas turbine engine airfoilcomprising: providing at least two gas turbine engine airfoil compositepreform components, interlocking the airfoil composite preformcomponents with a first locking component so that mating faces of theairfoil composite preform components face each other, and inserting afiller material between the mating faces of the airfoil compositepreform components.
 2. The method of claim 1, wherein one or more of theairfoil composite preform components are provided in a partiallyrigidized state.
 3. The method of claim 1, wherein the airfoil compositepreform components comprise woven fiber.
 4. The method of claim 1,wherein the airfoil composite preform components comprise an airfoilincluding at least one airfoil connecting tab and an endwall includingat least one endwall connecting tab, and the interlocking comprisesinserting the first locking component into a through hole in the atleast one airfoil connecting tab and a through hole in the at least oneendwall connecting tab.
 5. The method of claim 1, wherein the airfoilcomposite preform components comprise an airfoil including at least oneairfoil connecting tab providing the mating face of the airfoil, and anendwall including at least one endwall connecting tab providing themating face of the endwall, and the inserting comprises inserting thefiller material between the mating face of the at least one airfoilconnecting tab and the mating face of the at least one endwallconnecting tab.
 6. The method of claim 1, wherein the airfoil compositepreform components comprise an airfoil including at least one airfoilconnecting tab and an endwall including at least one endwall connectingtab and defining an opening substantially sized to receive therethroughthe at least one airfoil connecting tab, wherein the providing comprisesinserting the at least one airfoil connecting tab of the airfoil intothe opening of the endwall so that the at least one airfoil connectingtab is in side-by-side relation with the at least one endwall connectingtab.
 7. The method of claim 1, wherein the airfoil composite preformcomponents comprise an airfoil and an endwall, the airfoil including atleast one airfoil connecting tab and a seat portion, the endwalldefining an opening substantially sized to receive therethrough the atleast one airfoil connecting tab, wherein the providing comprisesinserting the at least one airfoil connecting tab of the airfoil intothe opening of the endwall so that the endwall rests over the seatportion.
 8. The method of claim 7, wherein the interlocking comprisesinserting the first locking component into a through hole in the atleast one airfoil connecting tab so that a rear portion of the firstlocking component rests over a portion of the endwall to interlock theendwall between the first locking component and the seat portion.
 9. Themethod of claim 7, wherein the seat portion provides the mating face ofthe airfoil and the portion of the endwall over the seat portionprovides the mating face of the endwall, and the inserting the fillermaterial comprises inserting the filler material between the mating faceof the airfoil and the mating face of the endwall.
 10. The method ofclaim 1, further comprising rigidizing the assembly of the airfoilcomposite preform components, the first locking component, and thefiller material.
 11. The method of claim 10 wherein the rigidizingcomprises using vapor infiltration.
 12. A method for forming a gasturbine engine airfoil assembly, comprising: providing at least two gasturbine engine airfoil composite preform components, interlocking theairfoil composite preform components with a first locking component, andinterlocking the first locking component and at least one of the airfoilcomposite preform components with a second locking component.
 13. Themethod of claim 12, wherein the airfoil composite preform componentscomprise an airfoil and an endwall, and the first locking componentinterlocks the airfoil and the endwall and the second locking componentinterlocks the first locking component and the endwall.
 14. The methodof claim 13, wherein the second interlocking comprises inserting thesecond locking component into a through hole in the first lockingcomponent and a through hole in the endwall.
 15. The method of claim 12,further comprising rigidizing the assembly of aircraft component preformcomponents, the first locking component, and the second lockingcomponent.
 16. The method of claim 15 wherein the rigidizing comprisesusing vapor infiltration.
 17. A gas turbine engine airfoil assembly,comprising: a composite airfoil including a pressure side and a suctionside having a hollow therebetween, and an airfoil connecting tabextending from the pressure side in a spanwise direction and an airfoilconnecting tab extending from the suction side in the spanwisedirection, and a composite endwall transverse to the composite airfoiland including a pressure side endwall connecting tab extending in thespanwise direction and a suction side endwall connecting tab extendingin the spanwise direction, pressure side locking means that prevent orinhibit relative lateral movement between the pressure side endwallconnecting tab and the pressure side airfoil connecting tab, and suctionside locking means that prevent or inhibit relative lateral movementbetween the suction side endwall connecting tab and the pressure sideairfoil connecting tab.
 18. The gas turbine engine airfoil assembly ofclaim 17 wherein the pressure side locking means comprises pressure sidelocking members that project transversely through the pressure sideendwall connecting tab and the pressure side airfoil connecting tab, andthe suction side locking means comprises suction side locking membersthat project transversely through the suction side endwall connectingtab and the pressure side airfoil connecting tab.
 19. The gas turbineengine airfoil assembly of claim 18 wherein the pressure side lockingmembers and suction side locking members comprise locking pins.
 20. Thegas turbine engine airfoil assembly of claim 17 comprising an integralcomposite structure.